Machinable CMC insert

ABSTRACT

An assembly comprising a ceramic matrix composite component, a monolithic insert, and a polymer char and a method for producing the same. The ceramic matrix composite component may include an exterior surface, with the monolithic insert bonded to the exterior surface of the ceramic matrix composite component. The polymer char may be sandwiched between the monolithic insert and the exterior surface of the ceramic matrix composite that may bond the monolithic insert to the ceramic matrix composite component.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority to and the benefit of U.S. ProvisionalPatent Application No. 62/171,280, filed 5 Jun. 2015, the disclosure ofwhich is now expressly incorporated herein by reference.

FIELD OF THE DISCLOSURE

The present disclosure relates generally to ceramic matrix compositecomponents, and more specifically to machinable inserts.

BACKGROUND

Gas turbine engine components are exposed to high temperatureenvironments with an increasing demand for even higher temperatures.Economic and environmental concerns relating to the reduction ofemissions and the increase of efficiency are driving the demand forhigher gas turbine operating temperatures. In order to meet thesedemands, temperature capability of the components in hot sections suchas blades, vanes, blade tracks, seal segments and combustor liners mustbe increased.

Ceramic matrix composites (CMCs) may be a candidate for inclusion in thehot sections where higher gas turbine engine operating temperatures arerequired. One benefit of CMC engine components is the high-temperaturemechanical, physical, and chemical properties of the CMCs which allowthe gas turbine engines to operate at higher temperatures than certaincurrent engines.

SUMMARY

The present disclosure may comprise one or more of the followingfeatures and combinations thereof.

According to an aspect of the present disclosure, an assembly for use ina gas turbine engine is taught. The assembly may comprise a ceramicmatrix composite, an insert, and a polymer char. The ceramic matrixcomponent may have an exterior surface. The insert may be bonded to theexterior surface of the ceramic matrix composite component. The polymerchar may be sandwiched between the monolithic insert and the exteriorsurface of the ceramic matrix composite component to bond the monolithicinsert to the ceramic matrix composite component.

According to another aspect of the present disclosure, a method ofjoining an insert to a ceramic matrix composite component for use in agas turbine engine is taught. The method may comprise providing aceramic preform comprising silicon carbide fibers, depositing apre-ceramic polymer along an exterior surface of the ceramic preform,positioning the insert along the exterior surface of the ceramic preformsuch that the pre-ceramic polymer is sandwiched between the ceramicpreform and the insert, and heating the pre-ceramic polymer to form apolymer char that bonds the insert to the ceramic preform.

These and other features of the present disclosure will become moreapparent from the following description of the illustrative embodiments.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cut-away perspective view of a gas turbine engine showingthat the gas turbine engine includes a compressor section, a combustorsection, and a turbine section that cooperate to drive an output shaft;

FIG. 2 is a schematic of a structural component, a metallic component,and an insert sandwiched between the structural component and themetallic component;

FIG. 3 is a schematic of a structural component, a metallic component,and an insert including a polymer char sandwiched between the insert andthe structural component;

FIG. 4 is a schematic view of a woven fibers of a structural componentand the chopped fibers of the insert;

FIG. 5 is a diagrammatic view of an insert extending entirely across theexterior surface of a blade track segment;

FIG. 6 is a detail view of an insert adapted to mate with a recess alongexterior surface of a blade track segment, the insert including aprotrusion for interlocking with the recess of the blade track segment;

FIG. 7 is a cross-sectional view of the turbine shroud of FIG. 1 showingthe ceramic blade track held in place by a metallic carrier;

FIG. 8 is a detail view of an insert adapted to include a positioningtab for mating the insert with a recess along the exterior surface ofthe blade track segment;

FIG. 9 is a detail view of an insert adapted to include a dovetail postarranged to mate with a recess along the exterior surface of the bladetrack segment;

FIG. 10 is a detail view of multiple inserts arranged along the exteriorsurface of the blade track;

FIG. 11 is a detail view of an insert and a portion of the blade tracksubstantially covered by a ply such that the insert sandwiched betweenthe ply and the blade track;

FIG. 12 is a diagrammatic view of a vane with an insert or sacrificiallayer on the exterior surface of the vane for machining;

FIG. 13 is a diagrammatic view of a blade including an airfoil extendingbetween an outer and inner platform including inserts or a sacrificiallayer for machining;

FIG. 14 is a diagrammatic view of a combustor tile adapted to include aplurality of inserts and attachment features;

FIG. 15 is a diagrammatic view of an aerodynamic exhaust tail with aninner wall and an outer wall separated by a plurality of machinableinserts;

FIG. 16 is a block diagram depicting the method of joining the insert tothe structural component including a pre-ceramic polymer;

FIG. 17 is a block diagram depicting the method of joining the insert tothe structural component including co-infiltrating the insert and thestructural component;

FIG. 18 is a block diagram depicting the method of joining the insertand the structural component including covering a portion of the insertand the structural component with a ply; and

FIG. 19 is a block diagram depicting the method of joining thesacrificial layer or insert with the structural component.

DETAILED DESCRIPTION OF THE DRAWINGS

For the purposes of promoting an understanding of the principles of thedisclosure, reference will now be made to a number of illustrativeembodiments illustrated in the drawings and specific language will beused to describe the same.

As shown in FIG. 1, an illustrative aerospace gas turbine engine 10 mayinclude an output shaft 12, a compressor section 14, a combustor section16, and a turbine section 18 mounted to a case 20. The output shaft 12may be coupled to a propeller (not shown) and may be driven by theturbine section 18. The compressor section 14 may compress and deliverair to the combustor section 16. The combustor section 16 may mix fuelwith the compressed air received from the compressor section 14 toignite the fuel. The hot high pressure products of the combustionreaction in the combustor section 16 may be directed into the turbinesection 18 and the turbine section 18 may extract work to drive thecompressor section 14 and the output shaft 12 as suggested in FIG. 1.

The hot sections of the gas turbine engine 10 may benefit from the useof CMC components. CMC components may allow for higher operatingtemperatures and greater efficiencies. CMC components may need to bemachined to fit the tight tolerance requirements. The ability to meetthe tight tolerance requirements may allow reduced thickness of coatingsand abradable coatings that would otherwise be needed to achieve thetight tolerance requirements. Machining of the CMC component may lead toenvironmental attack of the CMC component. Machining of a CMC componentmay lead to cut fibers. The cut fibers may cause the fibers to beexposed to the environment. In some instances, the cut fibers may causethe CMC component to have a lower tolerance when compared a CMCcomponent with unexposed, uncut fibers. The machining and exposure ofthe fibers may result in cracks throughout the CMC component.

An illustrative assembly 10 for use in a gas turbine engine 10 may allowfor machining a CMC component without attacking the fibers. As shown inFIG. 2, the assembly 10 may include a structural component 24 having anexterior surface 26, a metallic component 30 spaced apart from thestructural component 24, and a sacrificial layer or insert 28 bonded tothe exterior surface 26 of the structural layer 24 and sandwichedbetween the metallic component 30 and the structural component 24. Thestructural component 24 described herein may include a CMC component.The structural component 24 may form a blade track, a vane, a blade, acombustor, a combustor tile, a faring, an exhaust tail cone, an exhaustliner flap, or the like for use in a gas turbine engine 10. The insert28 may provide a machinable surface to reduce machining of the CMCstructural component 24 during the machining process. In the absence ofinsert 28, machining of the structural component 24 may result inexposed fibers in the structural component 24.

The structural component 24 may be substantially homogenous and mayinclude Si-containing ceramic such as silicon carbide (SiC) or siliconnitride (Si₃N₄); boron carbide (B₄C), zirconium diboride (ZrB₂),molybdenum carbide (Mo₂C) or a similar silicon containing material. Inother examples, Structural component 24 may include a metal silicide,such as a molybdenum-silicon alloy (e.g., MoSi₂) or a niobium-siliconalloy (e.g., NbSi₂). The structural component 24 may include a matrixmaterial and a reinforcement material. The matrix material may include aceramic material such as SiC, Si₃N₄, B₄C, ZrB₂ Mo₂C or the like. In someexamples, the reinforcement material may include a continuousmonofilament or multifilament weave. The reinforcement material mayinclude SiC, Si₃N_(4;) or the like.

The structural component 24 may include fibers as described above, whichmay be coated with boron nitride, pyrolytic carbon, oxide interfacecoating, or the like. The structural component 24 may be a 2D laminate,a 3D weave, or any other composite structure.

As shown in FIG. 2, the structural component 24 includes an exteriorsurface 26. The exterior surface 26 of the structural component 24 maybe substantially bonded to the insert 28. The exterior surface 26 may bea chemically homogenous surface for bonding the insert 28. The exteriorsurface 26 may include various recesses and protrusions for holding,bonding, or joining with the insert 24. The geometry of the exteriorsurface 26 may be determined based on the use of the structuralcomponent 24 within the gas turbine engine 10. As shown in FIG. 9 theinsert 28 may cover the entire area of the structural component 24. Asshown in FIGS. 6-7 and 10-12 the insert 28 may cover only a portion ofthe structural component 24.

The insert 28 may be bonded to the exterior surface 26 of the structuralcomponent 24 and may be sandwiched between the CMC component 24 and themetallic component 30. The insert 28 may be bonded to the exteriorsurface of the structural component 24 to prevent degradation andcutting of the fibers of the structural component 24 during machiningthe assembly to the final specifications for use in a gas turbineengine. Machining of the ceramic fibers of the structural component 24may result in cracks and reduced tolerance requirements, so machiningthe insert 28 instead of the structural component 24 may be beneficial.

The insert 28 may include ceramic materials, powder, or resin char. Theceramic materials of the insert 28 may include chopped carbon fibers,chopped silicon carbide fibers, or the like. The insert 28 may bebetween about 0.005 inches thick and about 0.04 inches thick dependingon the location of the insert 28 within the gas turbine engine 10.

As shown in FIG. 3, in some examples the assembly 20 may further includea polymer char 32 or resin char sandwiched between the insert 28 and thestructural component 24. The polymer char 32 may be pre-ceramic polymerresin char. A polymer char may be a solid ceramic material that formswhen the pre-ceramic polymer is heat treated at an elevated temperature.The polymer char 32 may be joined using an adhesive to a pre-ceramiccomponent and infiltrated, as described below, to bond the insert 28 tothe structural component. The polymer char 32 may form a continuousceramic matrix between the structural component 24 and the insert 28such that the resin char may substantially bond or join the structuralcomponent 24 and the insert 28. The polymer char 32 may include silicon,carbon, silicon carbide, a binding agent, oxycarbide silicon oxynitride,silicon nitride, or the like. The polymer char 32 may include ceramicfibers. The ceramic fibers may include chopped fibers, woven fibers,unwoven fibers, or the like.

In some examples, the insert 28 may be a resin char insert. The resinchar insert may be formed from a polymer char, such as the polymer char32 described above. The resin char insert may be substantially formedfrom the polymer char 32 to bond the insert 28 to the structuralcomponent 24, according to the method described below. The resin charinsert may form a continuous ceramic matrix between the structuralcomponent 24 and the resin char insert subsequent to infiltration andheating of the resin char to form the polymer char as described below.The polymer char 32 may include silicon, carbon, silicon carbide, abinding agent, oxycarbide silicon oxynitride, silicon nitride, or thelike. The polymer char 32 may include ceramic fibers. The ceramic fibersmay include chopped fibers, woven fibers, unwoven fibers, or the like.

In some examples, the insert 28 may include silicon, silicon carbide.The structural component 24 and the insert 28 may be placed in a preformtool prior to infiltration with silicon metal, silicon alloy or thelike. The structural component 24 and the insert 28 may beco-infiltrated with silicon metal, silicon alloy, or the like to formthe continuous, uninterrupted silicon carbide matrix between thestructural component 24 and the insert 28 to bond or join the structuralcomponent 24 and the insert 28.

The insert 28 or sacrificial layer may comprise ceramic fibers. Theceramic fibers may include silicon carbide fibers, silicon fibers, orthe like. The ceramic fibers of the insert 28 may be unarranged fibers.Unarranged fibers may be unwoven, loosely braided, chopped fibers, orthe like. The ceramic fibers of the insert 28 may be substantially wovenfibers which may remain woven, or may be chopped after weaving of thefibers.

As shown in FIG. 11, in some examples, the insert 28 may be partially orsubstantially covered by a ply 34 such that the insert 28 is sandwichedbetween the ply 34 and the structural component 24. The ply 34 mayextend beyond the insert 28 in at least one direction so that the ply 34is joined to the insert 28 and the structural component 24. The ply 34may be at least one layer of ceramic fibers. The ply 34 of fabric may beapplied to the insert 28 and the structural component 24 prior to slurryinfiltration, such that the ply 34 is further bonded to the insert 28and structural component 24 after infiltration. The ply 34 may be bondedto the insert 28 and the structural component 24 by an organic tackyagent such as PVA may assist with adhering the ply to stick to theinsert 28 to allow for infiltration of the ply 34, the inset 28, and thestructural component 24. The infiltration of the silicon metal may forma silicon carbide matrix which may extend from the structural component24 to the ply 34 to bond the Structural component 24, the insert 28, andthe ply 34 together.

In some examples, the ply 34 may include a layer of woven siliconcarbide fiber. A second layer of silicon carbide fiber may be placed onthe first silicon carbide fiber to form a weave or fabric of siliconcarbide fibers. Any suitable number of layers of silicon carbide fibersmay be used to provide the desired protection to the insert 28 and thestructural component 24. The ply 34 or an exterior layer of the ply 34may be locally machined away instead of machining of the fibers of thestructural component 24. In some examples, an additional ply may beplaced between the insert 28 and the structural component 24 to assistwith bonding or joining the insert 28 and the structural component 24.

In some examples, the ceramic fibers may be coated with boron nitride, aCVD pyrolytic carbon coating, a silicon doped boron nitride coating, orthe like. In other examples, the ceramic fibers may be substantiallyuncoated. The substantially uncoated fibers may not undergo the CVIprocess infiltration process as described below and may be bare siliconcarbide fibers. Fibers without the boron nitride coating may be moreeasily machined and may provide less environmental attack on the insert28 or the structural component 24.

In some examples, the volume fiber fraction of the insert 28 may belower than the volume fiber fraction of the structural component 24. Thevolume fiber fraction may be the volume of fibers as a fraction of thetotal volume of the component. The lower volume fiber fraction of theinsert 28 may allow for fewer fibers to be machined away during themachining process to prevent cracking and reduction in tolerance of thestructural component 24. The lower volume fiber fraction of the insert28 may also allow for improved infiltration of the structural component24.

In some examples, the insert 28 may include a powder and/or a bindingagent. The powder to form the insert 28 may include silicon carbide,silicon, or any ceramic containing powder. The powder may be a loosepowder or a pressed powder. The pressed powder may be pressed into acompact of the final shape of the insert. The powder may be infiltratedwith silicon metal to produce a silicon carbide-silicon carbide matrixthroughout both the insert 28 and the structural component 24. Thepowder may include a binding agent such as a polymeric binder to assistwith binding the insert 28 to the structural component 24. In someexamples, the insert 28 may be a powder referred to as a green bodyceramic. A green body ceramic may be an un-infiltrated ceramic componentincluding loose or compact silicon carbide powder, which may beinfiltrated to form a silicon carbide-silicon carbide matrix.

In some examples, the insert 28 may include reticulated foam or amaterial of substantially continuous porosity. Continuous porosity maybe a permeable structure with open cells for infiltrating materials. Theinsert 28 may be a ceramic foam including silicon, silicon carbide, orthe like. The reticulated foam may have a porosity of between about 10%and about 90% by volume such that it may be infiltrated in subsequentdensification steps to bond the insert 28 to the structural component24. The continuous porosity may allow gas or liquid phase silicon toinfiltrate into the pores during infiltration steps. The reticulatedfoam may be machinable such that fibers of the structural component 24are not exposed during the machining step.

As shown in FIG. 4, the insert 28 may include a needle punched layer.The needle punched layer 29 may include woven fibers 31, unwoven fibers29, or a combination of woven and unwoven fibers. The needle may punchthrough the layer of fibers to provide chopped fibers. The fibers may becoated with boron nitride, or the fibers may be substantially uncoatedprior to needle punching. The needle punched insert may have a lowerfiber fraction volume to allow for improved infiltration of thestructural component 24. In some examples, the insert 28 may includechopped fiber. As shown in FIGS. 4 and 10, at least one insert 28 may bepositioned along the exterior surface of the blade track segment 40. Theinsert 28 may include chopped fibers. The fibers may be chopped by anynumber of methods including needle punching as described below. As shownin FIG. 4, the blade track segment 40 may include woven fibers. Thewoven fibers may be arranged in any number of weaves including a 0degree, 90 degree weave, a five point satin harness, a 7 point satinharness or any of the weaves described herein. The chopped fiber insert28 may include a lower volume fiber fraction compared to the blade tracksegment 40. In several of the examples below the structural component 24is a blade track segment 40.

As shown in FIG. 5, the insert 28 may extend entirely across in at leastone direction of the exterior surface 26 of the structural component 24such as the blade track segment 40. As shown in FIG. 6, in someembodiments, insert 28 may cover only a portion of the exterior surface26 of the blade track segment 40. The desired area of coverage for theinsert may be determined by the machining and tolerance requirements ofthe final component.

As shown in FIG. 7, the insert 28 may be sandwiched between a carrier 38and a blade track segment (sometimes called a seal ring) 40. The carrier38 may be an annular, metallic component and may support the blade tracksegment 36 in position adjacent to the blades of the turbine wheelassembly. The blade track segment 40 may include a runner 41, a forwardattachment arm 45 and an aft attachment arm 47 as shown. The runner 41may extend around a turbine wheel assembly to block gasses from passingover the turbine blades without pushing the blades 49. The forwardattachment arm 45 may have a radially-extending portion 51 and may havean axially-extending portion 53. The aft attachment arm 47 may have aradially-extending portion 55 and an axially-extending portion 57 forattaching to the carrier 32. The blade track segment 40 may includesilicon containing ceramic fibers and a reinforcement matrix asdescribed herein. The blade track 41 may include or be formed of asilicon-carbide/silicon-carbide ceramic matrix composite The insert 28may be sandwiched between the carrier 38 and the blade track 40 in anynumber of configurations with examples described below.

As shown in FIG. 6, the insert 28 and the blade track segment 40 may bejoined by a recess 44 along the exterior surface 26 of the blade tracksegment 40. The recess 44 may be adapted to include any geometry forholding the insert 28. The insert 28 may include a protrusion 46 formating with the recess 44 of the blade track segment 40. The protrusion46 may be formed of the same material as the insert 28 such that theprotrusion 46 is an extension of the insert 28.

As shown in FIG. 8, in some embodiments the insert 28 may include apositioning tab 42 for mating the insert 28 with a recess 44 of theblade track segment 40. The positioning tab 42 may assist with “locking”the insert 28 in the blade track segment 40 such that the componentswill not become separated during use of the assembly. The positioningtab 42 may assist with reducing axially movement of the insert 28 afterinfiltration and processing as described herein. In some examples, thepositioning tab 42 may assist with locating the insert 28 along theblade track segment 40. The recess 44 of the of the blade track segment40 may be made in the blade track segment 40 when the fibers arearranged and laid up to form the ceramic preform component describedabove.

As shown in FIG. 9, in some examples the insert may include a dovetailpost 43 arranged to mate with the recess 44 of the blade track segment40. The dovetail post 43 may allow the insert 28 to be locked into theblade track segment 40 to prevent movement of the insert 28. Asdescribed below with regard to FIG. 11, an overwrap or ply 34 may not beincluded when the insert 28 includes the dovetail post 43. The recess 44of the of the blade track segment 40 for holding the insert 28 may bemade in the blade track segment 34 when the fibers are arranged and laidup to form the ceramic preform component described above.

As shown in FIG. 10, in some examples a plurality of inserts 28 may beused. A plurality of inserts 28 may be used when a plurality oflocations along an exterior surface of a blade track 40 may requiremachining. The inserts 28 may be placed in locations which may requiremachining to achieve the necessary tolerance requirements for the bladetrack 40. The inserts 28 may be placed at any location along the bladetrack 40 or component.

As shown in FIG. 11, in some examples, the insert 28 and a portion ofthe blade track 40 may be partially or substantially covered by a ply 34such that the insert 28 is sandwiched between the ply 34 and the bladetrack 40. The ply 34 may extend beyond the insert 28 in at least onedirection so that the ply 34 is joined to the insert 28 and thestructural component blade track 40. The ply 34 may include at least onelayer of ceramic fibers as described above.

In some examples, the structural component 24 may include an airfoilsuch as a blade 50 as shown in FIG. 13 or a vane 58 as shown in FIG. 12.The blade 58 adapted for use in a turbine section of a gas turbineengine includes an airfoil 52. The airfoil 52 extends between outer andinner platforms 54, 56 of the blade 50. Although only one airfoil 52 isshown to extend between the platforms 54, 56 in FIG. 13, a plurality ofairfoils 52 may extend between the annular platforms 54, 56. Theplurality of airfoils 52 are circumferentially spaced such that theairfoils 52 and the platforms 54, 56 cooperate to direct fluid flowingthrough the turbine section toward downstream sections of the gasturbine engine. The insert 28, or sacrificial layer described above, maybe positioned along an exterior surface of the vane to provide asacrificial layer for machining. The sacrificial layer for machining mayallow the airfoil structure to be tightly machined to fit thegeometrical tolerances.

As shown in FIG. 14, in some examples, the structural component 24 mayinclude a combustor tile 60. The combustor tile 60 may be constructed ofa ceramic matrix composite material. The combustor tiles 60 may bearranged around the circumference of an outer or inner shell of thecombustor. The combustor tile 60 may include a plurality of inserts 28and attachment features 62 to space the combustor tile 60 off of anexterior full-hoop liner. The attachments 62, shown in FIG. 14, aredepicted as fir-tree attachments, but dovetail attachments or any othersuitable attachment may be used. The inserts 28 and attachment features62 may be processed with the combustor tile 60, as described below, andmay then be machined to become retention features. The inserts 28 may beused to position or locate the combustor tile 60 off of an outerfull-hoop liner.

As shown in FIG. 15, in some examples the structural component 24 mayinclude an exhaust cone 70. The insert 28 may be placed between an outerwall 80 and an inner wall 82, where the outer wall 80 and the inner wall82 may have differing coefficients of thermal expansion. The insert 28may be machined to fit between the two walls 80, 82 to allow for agreater tolerance of thermal expansion than would otherwise be permittedwith two differing materials.

An illustrative method for joining the insert to the structuralcomponent described herein may include providing a ceramic preformcomprising silicon carbide fibers. The ceramic preform may form thestructural component 24 according to the methods described below.

As shown in FIG. 16, an illustrative method 110 for joining an insert toa structural component is provided. According to a step 112 of themethod 110, a ceramic preform comprising silicon carbide fibers isprovided. The ceramic preform may be produced according to any of themethods described below.

In a step 114 of the method 110, a pre-ceramic polymer may be depositedalong an exterior surface of the ceramic preform. The pre-ceramicpolymer may be added as an adhesive, which may form the char after heattreating. The pre-ceramic polymer may be pre-ceramic polymer resin char.The pre-ceramic polymer may form a solid ceramic material when thepre-ceramic polymer is heated to an elevated temperature. In someembodiments, the polymer char may be joined to the structural componentpreform using an adhesive. The adhesive may bond the insert to thestructural component for infiltration. The polymer char may includepre-ceramic phases, silicon carbide, transition metals, transition metalborides, transition metal silicides or combinations thereof. In someexamples, the pre-ceramic polymer may include carbon-based polymersystems such as phenolic resin and furfuryl alcohol resin. The char mayhave similar chemical properties to the pre-ceramic polymer with thesome of the chemical elements removed by the heating process. Theelements which may be removed may include hydrogen, oxygen, andnitrogen. Specifically the pre-ceramic material may include SMP-10™, asilicon carbide matrix precursor sold by Starfire Systems, or otherprecursors with similar properties to SMP-10. In a step 116 of themethod 110, the insert is positioned along the exterior surface of theceramic preform. The insert is positioned such that the pre-ceramicpolymer is sandwiched between the ceramic preform and the insert. In astep 118 of the method 110 the pre-ceramic polymer is heated to form apolymer char that bonds the insert to the ceramic preform.

The insert may be co-infiltrated with the structural component preformusing chemical vapor infiltration, chemical vapor deposition, slurryinfiltration, melt infiltration, polymer impregnation and pyrolysis orany combination as described below. The pre-ceramic polymer may beheated in a furnace and/or may be heated through the infiltrationprocesses. Heating of the pre-ceramic polymer may be performed throughthe processes of CVI, SMI, or brazing. As the pre-ceramic polymer isheated the pre-ceramic polymer may form a ceramic matrix which mayextend to the structural component preform to join the insert and thestructural component. The pre-ceramic polymer char may be heated to atemperature between about 1300° C. and about 1500° C. to form thepolymer resin char.

As shown in FIG. 17, an illustrative method for joining an insert to aceramic matrix composite structural component is described in method210. According to a step 212 of the method 210, a ceramic preformcomprising silicon carbide fibers is provided. The ceramic preformcomprising silicon carbide fibers may be provided according to themethods described above.

In a step 214 of the method 210, the ceramic insert is positionedadjacent to the ceramic preform. The ceramic insert may be positionedsuch that the insert and the preform are held together in a tool priorto infiltration. In some embodiments the insert may be positioned withina recess along the surface of the ceramic preform. In some embodiments,the ceramic insert may be adhered to the ceramic preform using anadhesive to join the ceramic insert and ceramic preform prior toinfiltration.

In a step 216 of the method 210, the insert and the ceramic preform maybe co-infiltrated with a silicon metal or silicon alloy to form asilicon carbide matrix extending from within the ceramic preform towithin the insert thereby joining the insert to the ceramic preform. Thestep of co-infiltrating the ceramic preform and the insert may includechemical vapor infiltration, chemical vapor deposition, slurryinfiltration, melt infiltration, polymer impregnation and pyrolysis, ora combination thereof. The steps of infiltrating with silicon metal orsilicon alloy are described above.

In some embodiments, the insert may comprise a powder such as silicon,silicon carbide, or a combination thereof. The powder may be pressed toform the insert. The powder may be pressed into a compact via eithercold pressing or hot pressing of the powder. In some embodiments, apolymeric binder may be added to the powder to assist with formation ofthe insert.

As shown in FIG. 18, an illustrative method 310 for joining an insert toa ceramic matrix composite structural component is provided. Accordingto a step 312 of the method 310, a ceramic preform comprising siliconcarbide fibers is provided. The ceramic preform comprising siliconcarbide fibers may be produced according to any of the methods describedabove.

In a step 314 of the method 310, a ceramic insert is positioned adjacentto the ceramic preform. The ceramic insert may be positioned such thatthe insert and the preform are held together in a tool prior toinfiltration. In some embodiments, the insert may be positioned within arecess along the surface of the ceramic preform. The ceramic insert maybe adhered to the ceramic preform using an adhesive or tacky agent tojoin the ceramic insert and ceramic preform prior to infiltration.

In a step 316 of the method 310, at least a portion of the insert and atleast a portion of the ceramic preform may be covered with a ply. Theply may be a fabric laid into the CVI tool. The insert may then be laidon top of the within the tool such that the insert may be sandwichedbetween the ply and the ceramic preform prior to infiltration. The plymay comprise at least one layer of silicon carbide fibers. The ply maybe any suitable number of silicon carbide fibers to achieve thethickness desired for the insert. In a step 318 of the method 310, theceramic preform and the ply may be co-infiltrated. Co-infiltrating maybe performed according to the methods described above. Co-infiltratingof the ply, the insert and the structural component may provide morecomplete consolidation of the component.

As shown in FIG. 19, an illustrative method 410 for joining an insert toa ceramic matrix composite structural component is provided. Accordingto a step 412 of the method 410, a ceramic preform comprising siliconcarbide fibers is provided. The ceramic preform comprising siliconcarbide fibers may be produced according to the methods described above.

In a step 414 of the method 410, the structural component and the insertor sacrificial layer may be infiltrated with silicon metal or siliconalloy to join the insert and the structural component. The infiltrationmay be performed via chemical vapor infiltration, chemical vapordeposition, slurry infiltration, melt infiltration, polymer impregnationand pyrolysis, or a combination thereof. The steps of infiltrating withsilicon metal or silicon alloy are described above. The infiltrationwith silicon metal or silicon alloy may produce a ceramic matrix whichmay extend form the structural component to the insert or sacrificiallayer.

The structural component and the insert may be formed according to themethods described below. Chemical vapor deposition (CVD) or chemicalvapor infiltration (CVI) (CVD and CVI collectively referred herein asCVI) may be used to build up one or more layers on the ceramic fibers ofthe structural component preform. The one or more layers may include asilicon carbide layer. Furthermore, one or more intermediate layers suchas boron nitride may be deposited prior to the silicon carbide layer.CVD may follow the same thermodynamics and chemistry. CVI and CVD may benon-line of sight processes process such that deposition can occur onthe ceramic fibers that are within or internal to the preform.Furthermore, such CVI and CVD may occur completely within a furnace. Thestarting material for CVI may include a gaseous precursor controlled byquartz tubes and may be performed at temperatures between about 900° C.and about 1300° C. CVI may be performed at a relatively low pressure andmay use multiple cycles in the furnace. Silicon carbide may also bedeposited to build up one or more layers on the fibers while the preformis in the furnace. The silicon carbide may provide additional protectionto the fibers and may also increase the stiffness of the structuralcomponent preform. In some examples, boron nitride may be depositedprior the silicon carbide to provide further beneficial mechanicalproperties to the fibers. The preform may be taken out of the furnaceafter a deposition and weighed. If the preform is not at the targetweight it may go through the furnace for another run, which may occur asmany times as necessary in order to achieve the target weight. Thetarget weight may be determined by the final part to be made. CVI mayform a preform with a porosity of between about 40% and about 50%. Ifthe preform is at the target weight, the part may undergo slurryinfiltration.

Once the structural component preform fibers are coated via CVI,additional particles may be infiltrated into the preform via otherinfiltration methods. For example, a slurry infiltration process mayinclude infiltrating the preform with slurry. Dispersing the slurrythroughout the preform may include immersing the preform in the slurrycomposition. The slurry may include particles of carbon and/or siliconcarbide. The slurry may flow into the spaces, pores, or openings betweenthe fibers of the preform such that the slurry particles may uniformlyimpregnate the pores of the preform and reside in the intersticesbetween the preform fibers. The slurry infiltration process may form apreform with a porosity of between about 35% and about 45%.

Prior to immersion, the preform fibers may optionally be prepared forslurry infiltration by exposing the fibers to a solution including, forexample, water, solvents, surfactants and the like to aid impregnationof the fibers. Optionally, a vacuum may be drawn prior to slurryintroduction to purge gas from the preforms and further enhanceimpregnation. Slurry infiltration may be conducted at any suitabletemperature such as at room temperature (about 20° C. to about 35° C.).The slurry infiltration may be enhanced by application of externalpressure after slurry introduction such as at one atmosphere pressuregradient.

After slurry infiltration, the structural component preform may undergomelt infiltration. During melt infiltration a molten metal or alloy maywick between the openings of the preforms. In various embodiments, themolten metal or alloy may have a composition that includes silicon,boron, aluminum, yttrium, titanium, zirconium, oxides thereof, andmixtures and combinations thereof. In some instances, graphite powdermay be added to assist the melt infiltration. The molten metal or alloymay wick into the remaining pores of the preform through capillarypressure. For example, molten silicon metal may wick into the pores andform silicon carbide to create a matrix between the fibers resulting ina relatively structural component. For example, structural component hasdensified, the structural component may have a porosity of between about1 percent and about 10 percent by volume. In one example, a temperatureof the molten silicon metal may be between about 1400° C. and about1500° C. for infiltration. The duration of the infiltration may bebetween about 15 minutes and 4 hours. The infiltration process may becarried out under vacuum, but in other embodiments melt infiltration maybe carried out with an inert gas under atmospheric pressure to limitevaporation losses.

In some embodiments, the insert or sacrificial layer may at leastpartially wrap around the structural component. The sacrificial layermay wrap entirely around the component or may only cover a portion ofthe structural component.

In some embodiments, ceramic fibers may be chopped to create the insertor sacrificial layer. The fibers may be chopped by using a needle punchmethod. The needle punch method may include feeding the fibers through aneedle loom wherein the needles punch through the fibers at least onetime. The needle punches through the layer of fibers multiple timesuntil the layer of fibers are chopped and the fibers may then be drawnoff the needle loom. The fibers may also be open braided or unwoven andmay not need to go through the needle punch method. Any suitable methodfor providing substantially unwoven, chopped, or unbraided fibers may beused. Needle punching may be performed prior to bonding the insert tothe structural layer or may be performed after bonding the insert andthe structural component. Needle punching may not go into the structuralcomponent.

In some embodiments, the insert may be rigidized prior to joining theinsert to the ceramic structural preform. The insert may be rigidizedvia chemical vapor deposition or chemical vapor infiltration similarlyto the method of rigidizing or infiltrating the preform described below.After CVI the insert may have a porosity of between about 0% by volumeand about 50% by volume. The partially rigidized, but still porousinsert may then be joined to the ceramic structural preform andrigidized or densified further through the melt infiltration and slurryinfiltration processes. In this embodiment, the insert and ceramicstructural preform may be co-infiltrated through slurry infiltration andmelt infiltration contemporaneously. After co-infiltration the insertmay have a porosity of between about 0% by volume and about 10% byvolume.

In some embodiments, the tool for holding the insert, the ply, thestructural component or any combination thereof may have a recess forpositioning the components. The recess in the tool may also help to formthe recess within the structural component for mating with the insert.

In some embodiments, after any combination of methods described hereinthe insert or sacrificial layer may be machined to a final shape.Machining of the insert may allow the structural component to achievethe necessary geometry and tolerance requirements without exposing thestructural component to the environment. The insert may be placed alongany surface of the structural component which may require machining.

While the disclosure has been illustrated and described in detail in theforegoing drawings and description, the same is to be considered asexemplary and not restrictive in character, it being understood thatonly illustrative embodiments thereof have been shown and described andthat all changes and modifications that come within the spirit of thedisclosure are desired to be protected.

What is claimed is:
 1. An assembly for use in a gas turbine engine, theassembly comprising a ceramic matrix composite component having anexterior surface, a monolithic insert bonded to the exterior surface ofthe ceramic matrix composite component, and a polymer char sandwichedbetween the monolithic insert and the exterior surface of the ceramicmatrix composite component that bonds the monolithic insert to theceramic matrix composite component, wherein the ceramic matrix compositecomponent and the monolithic insert are co-infiltrated with silicon,silicon alloy, silicon carbide, or a combination thereof.
 2. Theassembly of claim 1, wherein the monolithic insert comprises a resinchar comprising silicon containing ceramic material, chopped carbonfibers, chopped silicon carbide fibers, or a combination thereof.
 3. Theassembly of claim 1, wherein the monolithic insert comprises a pressedpowder comprising silicon carbide powder and a binding agent.
 4. Theassembly of claim 1, wherein the monolithic insert comprises a ceramicmaterial comprising silicon carbide.
 5. The assembly of claim 1, whereinthe monolithic insert comprises a reticulated foam having a continuousporosity.
 6. The assembly of claim 1, wherein the ceramic matrixcomposite component comprises a blade track, a vane, a blade, acombustor tile, an exhaust tail cone, or a combination thereof.
 7. Theassembly of claim 1, wherein the ceramic matrix composite componentcomprises a recess along the exterior surface for mating with themonolithic insert.
 8. The assembly of claim 7, wherein the ceramicmatrix composite component comprises a blade track segment with therecess along the exterior surface.
 9. The assembly of claim 8, whereinthe monolithic insert comprises a dovetail post arranged to mate withthe recess of the blade track segment.
 10. The assembly of claim 8,wherein the monolithic insert comprises a positioning tab for matingwith the recess of the blade track segment.
 11. The assembly of claim 1,further comprising a ply of fabric wrapped over at least a portion ofthe monolithic insert and the ceramic matrix composite.
 12. The assemblyof claim 11, wherein the ply of fabric comprises silicon-carbide fibers.13. A method of joining a monolithic insert to a ceramic matrixcomposite component for use in a gas turbine engine, the methodcomprising providing a ceramic preform comprising silicon carbidefibers, depositing a pre-ceramic polymer along an exterior surface ofthe ceramic preform, positioning the monolithic insert along theexterior surface of the ceramic preform such that the pre-ceramicpolymer is sandwiched between the ceramic preform and the monolithicinsert, heating the pre-ceramic polymer to form a polymer char thatbonds the monolithic insert to the ceramic preform, and co-infiltratingthe ceramic preform and the monolithic insert with liquid silicon,silicon alloy, silicon carbide, or a combination thereof.
 14. The methodof claim 13, wherein co-infiltrating comprises chemical vaporinfiltration, chemical vapor deposition, slurry infiltration, meltinfiltration, polymer impregnation and pyrolysis, or a combinationthereof.
 15. The method of claim 14, wherein the pre-ceramic polymercomprises a ceramic phase, SiC, a transition metal, a transition metalboride, a transition metal silicide, or combinations thereof.
 16. Themethod of claim 13, wherein heating the pre-ceramic polymer is performedat a temperature between about 1300° C. and about 1500° C.
 17. Themethod of claim 13, further comprising placing a ply of fabric to coverat least a portion of the monolithic insert and the ceramic preform. 18.The method of claim 13, further comprising machining the monolithicinsert after bonding the monolithic insert to the ceramic preform. 19.The method of claim 13, wherein heating the pre-ceramic polymer to formthe polymer char is performed contemporaneously with at least someportion of co-infiltrating the ceramic preform and the monolithicinsert.